Interior cooling circuits in turbine blades

ABSTRACT

An airfoil of a turbine rotor blade that includes a cooling configuration having a plurality of elongated flow passages for receiving and directing a coolant along a path through the airfoil. The cooling configuration may include: a central flow passage flanked to each side by near-wall flow passages that includes a pressure side near-wall flow passage and a suction side near-wall flow passage; a first port that fluidly connects the central flow passage to the pressure side near-wall flow passage; a second port that fluidly connects the central flow passage to the suction side near-wall flow passage; and impingement connectors that fluidly connect the central flow passage to a leading edge flow passage.

BACKGROUND OF THE INVENTION

This invention relates to turbine airfoils, and more particularly to hollow turbine airfoils, such as rotor or stator blades, having internal channels for passing fluids such as air to cool the airfoils.

Combustion or gas turbine engines (hereinafter “gas turbines”) include a compressor, a combustor, and a turbine. As is well known in the art, air compressed in the compressor is mixed with fuel and ignited in the combustor and then expanded through the turbine to produce power. The components within the turbine, particularly the circumferentially arrayed rotor and stator blades, are subjected to a hostile environment characterized by the extremely high temperatures and pressures of the combustion products that are expended therethrough. In order to withstand the repetitive thermal cycling as well as the extreme temperatures and mechanical stresses of this environment, the airfoils must have a robust structure and be actively cooled.

As will be appreciated, turbine rotor and stator blades often contain internal passageways or circuits that form a cooling system through which a coolant, typically air bled from the compressor, is circulated. Such cooling circuits are typically formed by internal ribs that provide the required structural support for the airfoil, and include multiple flow paths designed to maintain the airfoil within an acceptable temperature profile. The air passing through these cooling circuits often is vented through film cooling apertures formed on the leading edge, trailing edge, suction side, and pressure side of the airfoil.

It will be appreciated that the efficiency of gas turbines increases as firing temperatures rise. Because of this, there is a constant demand for technological advances that enable turbine blades to withstand ever higher temperatures. These advances sometimes include new materials that are capable of withstanding the higher temperatures, but just as often they involve improving the internal configuration of the airfoil so to enhance the blades structure and cooling capabilities. However, because the use of coolant decreases the efficiency of the engine, new arrangements that rely too heavily on increased levels of coolant usage merely trade one inefficiency for another. As a result, there continues to be demand for new airfoil designs that offer internal airfoil configurations and coolant circulation that improves coolant efficiency.

A consideration that further complicates design of internally cooled airfoils is the temperature differential that develops during operation between the airfoils internal and external structure. That is, because they are exposed to the hot gas path, the external walls of the airfoil typically reside at much higher temperatures during operation than many of the internal ribs, which, for example, may have coolant flowing through passageways defined to each side of them. In fact, a common airfoil configuration includes a “four-wall” arrangement in which lengthy inner ribs run parallel to the pressure and suction side outer walls. It is known that high cooling efficiency can be achieved by the near-wall flow passages that are formed in the four-wall arrangement, however, the outer walls experience a significantly greater level of thermal expansion than the inner walls. This imbalanced growth causes stress to develop at the points at which the inner ribs and outer walls connect, which may cause low cyclic fatigue that can shorten the life of the blade. As such, the development of airfoil structures that use coolant more efficiently while also reducing stress caused by imbalanced thermal expansion between internal and external regions remains a significant technological industry objection.

BRIEF DESCRIPTION OF THE INVENTION

The present application thus describes an airfoil having a leading edge, a trailing edge, an outboard tip, and an inboard end where the airfoil attaches to a root configured to couple the turbine blade to a disc. The airfoil may further include a cooling configuration comprising a plurality of elongated flow passages for receiving and directing a coolant along a path through the airfoil. The cooling configuration may include: a central flow passage flanked to each side by near-wall flow passages that includes a pressure side near-wall flow passage and a suction side near-wall flow passage; a first port that fluidly connects the central flow passage to an upstream portion of the pressure side near-wall flow passage; a second port that fluidly connects the central flow passage to an upstream portion of the suction side near-wall flow passage; a leading edge flow passage; and impingement connectors that fluidly connect the central flow passage to the leading edge flow passage.

These and other features of the present application will become apparent upon review of the following detailed description of the preferred embodiments when taken in conjunction with the drawings and the appended claims.

BRIEF DESCRIPTION OF THE DRAWINGS

These and other features of this invention will be more completely understood and appreciated by careful study of the following more detailed description of exemplary embodiments of the invention taken in conjunction with the accompanying drawings, in which:

FIG. 1 is a schematic representation of an exemplary turbine engine in which certain embodiments of the present application may be used;

FIG. 2 is a sectional view of the compressor section of the combustion turbine engine of FIG. 1;

FIG. 3 is a sectional view of the turbine section of the combustion turbine engine of FIG. 1;

FIG. 4 is a perspective view of a turbine rotor blade of the type in which embodiments of the present invention may be employed;

FIG. 5 is a side sectional view of a turbine rotor blade having an inner wall configuration according to conventional design; and

FIG. 6 is a cross-sectional view of the turbine rotor blade of FIG. 5.

DETAILED DESCRIPTION OF THE INVENTION

As an initial matter, in order to clearly describe the current invention it will become necessary to select certain terminology when referring to and describing relevant machine components within a gas turbine. When doing this, if possible, common industry terminology will be used and employed in a manner consistent with its accepted meaning. Unless otherwise stated, such terminology should be given a broad interpretation consistent with the context of the present application and the scope of the appended claims. Those of ordinary skill in the art will appreciate that often a particular component may be referred to using several different or overlapping terms. What may be described herein as being a single part may include and be referenced in another context as consisting of multiple components. Alternatively, what may be described herein as including multiple components may be referred to elsewhere as a single part. Accordingly, in understanding the scope of the present invention, attention should not only be paid to the terminology and description provided herein, but also to the structure, configuration, function, and/or usage of the component.

In addition, several descriptive terms may be used regularly herein, and it should prove helpful to define these terms at the onset of this section. These terms and their definitions, unless stated otherwise, are as follows. As used herein, “downstream” and “upstream” are terms that indicate a direction relative to the flow of a fluid, such as the working fluid through the turbine engine or, for example, the flow of air through the combustor or coolant through one of the turbine's component systems. The term “downstream” corresponds to the direction of flow of the fluid, and the term “upstream” refers to the direction opposite to the flow. The terms “forward” and “aft”, without any further specificity, refer to directions, with “forward” referring to the front or compressor end of the engine, and “aft” referring to the rearward or turbine end of the engine. It is often required to describe parts that are at differing radial positions with regard to a center axis. The term “radial” refers to movement or position perpendicular to an axis. In cases such as this, if a first component resides closer to the axis than a second component, it will be stated herein that the first component is “radially inward” or “inboard” of the second component. If, on the other hand, the first component resides further from the axis than the second component, it may be stated herein that the first component is “radially outward” or “outboard” of the second component. The term “axial” refers to movement or position parallel to an axis. Finally, the term “circumferential” refers to movement or position around an axis. It will be appreciated that such terms may be applied in relation to the center axis of the turbine.

By way of background, referring now to the figures, FIGS. 1 through 4 illustrate an exemplary combustion turbine engine in which embodiments of the present application may be used. It will be understood by those skilled in the art that the present invention is not limited to this type of usage. As stated, the present invention may be used in combustion turbine engines, such as the engines used in power generation and airplanes, steam turbine engines, and other types of rotary engines. The examples provided are not meant to be limiting to the type of the turbine engine.

FIG. 1 is a schematic representation of a combustion turbine engine 10. In general, combustion turbine engines operate by extracting energy from a pressurized flow of hot gas produced by the combustion of a fuel in a stream of compressed air. As illustrated in FIG. 1, combustion turbine engine 10 may be configured with an axial compressor 11 that is mechanically coupled by a common shaft or rotor to a downstream turbine section or turbine 13, and a combustor 12 positioned between the compressor 11 and the turbine 13.

FIG. 2 illustrates a view of an exemplary multi-staged axial compressor 11 that may be used in the combustion turbine engine of FIG. 1. As shown, the compressor 11 may include a plurality of stages. Each stage may include a row of compressor rotor blades 14 followed by a row of compressor stator blades 15. Thus, a first stage may include a row of compressor rotor blades 14, which rotate about a central shaft, followed by a row of compressor stator blades 15, which remain stationary during operation.

FIG. 3 illustrates a partial view of an exemplary turbine section or turbine 13 that may be used in the combustion turbine engine of FIG. 1. The turbine 13 may include a plurality of stages. Three exemplary stages are illustrated, but more or less stages may be present in the turbine 13. A first stage includes a plurality of turbine buckets or turbine rotor blades 16, which rotate about the shaft during operation, and a plurality of nozzles or turbine stator blades 17, which remain stationary during operation. The turbine stator blades 17 generally are circumferentially spaced one from the other and fixed about the axis of rotation. The turbine rotor blades 16 may be mounted on a turbine wheel (not shown) for rotation about the shaft (not shown). A second stage of the turbine 13 also is illustrated. The second stage similarly includes a plurality of circumferentially spaced turbine stator blades 17 followed by a plurality of circumferentially spaced turbine rotor blades 16, which are also mounted on a turbine wheel for rotation. A third stage also is illustrated, and similarly includes a plurality of turbine stator blades 17 and rotor blades 16. It will be appreciated that the turbine stator blades 17 and turbine rotor blades 16 lie in the hot gas path of the turbine 13. The direction of flow of the hot gases through the hot gas path is indicated by the arrow. As one of ordinary skill in the art will appreciate, the turbine 13 may have more, or in some cases less, stages than those that are illustrated in FIG. 3. Each additional stage may include a row of turbine stator blades 17 followed by a row of turbine rotor blades 16.

In one example of operation, the rotation of compressor rotor blades 14 within the axial compressor 11 may compress a flow of air. In the combustor 12, energy may be released when the compressed air is mixed with a fuel and ignited. The resulting flow of hot gases from the combustor 12, which may be referred to as the working fluid, is then directed over the turbine rotor blades 16, the flow of working fluid inducing the rotation of the turbine rotor blades 16 about the shaft. Thereby, the energy of the flow of working fluid is transformed into the mechanical energy of the rotating blades and, because of the connection between the rotor blades and the shaft, the rotating shaft. The mechanical energy of the shaft may then be used to drive the rotation of the compressor rotor blades 14, such that the necessary supply of compressed air is produced, and also, for example, a generator to produce electricity.

FIG. 4 is a perspective view of a turbine rotor blade 16 of the type in which embodiments of the present invention may be employed. The turbine rotor blade 16 includes a root 21 by which the rotor blade 16 attaches to a rotor disc. The root 21 may include a dovetail configured for mounting in a corresponding dovetail slot formed in the perimeter of the rotor disc. The root 21 may further include a shank that extends between the dovetail and a platform 24, which is disposed at the junction of the airfoil 25 and the root 21 and defines a portion of the inboard boundary of the flow path through the turbine 13. It will be appreciated that the airfoil 25 is the active component of the rotor blade 16 that intercepts the flow of working fluid and induces the rotor disc to rotate. While the blade of this example is a turbine rotor blade 16, it will be appreciated that the present invention also may be applied to other types of blades within the turbine engine 10, including turbine stator blades 17. It will be seen that the airfoil 25 of the rotor blade 16 includes a concave pressure side outer wall 26 and a circumferentially or laterally opposite convex suction side outer wall 27 extending axially between opposite leading and trailing edges 28, 29 respectively. The sidewalls 26 and 27 also extend in the radial direction from the platform 24 to an outboard tip 31.

FIGS. 5 and 6 provide exemplary embodiments of internal wall structures that define a cooling configuration according to the present invention. As indicated, the cooling configuration may include a plurality of elongated flow passages for receiving and directing a coolant through the airfoil 25. The cooling configuration may be positioned near the leading edge 28 of the airfoil 25. In a preferred embodiment, the several flow passages that are included in the cooling configuration of the present invention are positioned in the forward half of the airfoil 25.

In general, as illustrated in FIGS. 5 and 6, the cooling configuration of the present invention includes a central flow passage 40 that is flanked to each side by near-wall flow passages 43, 44. The near-wall flow passages include a pressure side near-wall flow passage 43 and a suction side near-wall flow passage 44. Positioned forward of the central flow passage 40, a leading edge flow passage 42 may be positioned in close proximity and parallel to the leading edge 28 of the airfoil 25. A port 46 may fluidly connect the central flow passage 40 to a downstream portion of the pressure side near-wall flow passage 43. Another port 46 may fluidly connects the central flow passage 40 to a downstream portion of the suction side near-wall flow passage 44. Finally, impingement connectors 48 may fluidly connect the central flow passage 40 to the leading edge flow passage 42.

It will be appreciated that, between the pressure side outer wall 26 and the suction side outer wall 27, the cooling configuration of the present invention provides for the laterally stacking of three flow passages: the pressure side near-wall flow passage 43, which is disposed adjacent to the pressure side outer wall 26; the suction side near-wall flow passage 44, which is disposed adjacent to the suction side outer wall 27; and the central flow passage 40, which is disposed between the pressure side and the suction side near-wall flow passages 43, 44. Given the ports 46 that connect the downstream portions of the pressure side near-wall flow passage 43 and a downstream portion of the suction side near-wall flow passage 44, it will be appreciated that the flow through the central flow passage 40 represents the combined flow of coolant from the two near-wall flow passages 43, 44.

A number of impingement connectors 48 fluidly connect the central flow passage 40 to the leading edge flow passage 42. As illustrated, the leading edge flow passage 42 is positioned in close proximity to the leading edge 28 of the airfoil 25. In preferred embodiments, the leading edge flow passage 42 extends radially outward in spaced relation to the leading edge 28 of the airfoil 25. At one end, the leading edge flow passage 42 is positioned near the inboard end of the airfoil 25. At the opposite end, the leading edge flow passage 42 is positioned near the outboard tip of the airfoil 25. The impingement connectors 48 are configured to allow coolant to pass from the central flow passage 40 to the leading edge flow passage 42, while also impinging the flow of coolant against an inner surface of the wall forming the leading edge 28 of the airfoil 25. It will be appreciated that the leading edge 28 of the airfoil 25 is a region that requires a significant level of coolant, and that impinging the flow of coolant in this manner enhances its effectiveness. In preferred embodiments, the central flow passage 40 may extend radially alongside the leading edge flow passage 42. The many impingement connectors 48 may be radially spaced between inboard and outboard ends of the leading edge flow passage 42 so that the flow of coolant is evenly applied.

The leading edge flow passage 42 may include a number of surface outlets 52. These may be configured to provide an outlet through which exhausted coolant is expelled from the airfoil 25. It will be appreciated that the surface outlets 52 also provide outlets through which film cooling may be applied to targeted surface areas of the airfoil 25.

In a preferred embodiment, the pressure side near-wall flow passage 43 includes axially-stacked and parallel first and second flow passages. As illustrated, each of these flow passages 43 may be defined on one side by the pressure side outer wall 26 of the airfoil 25. The pressure side near-wall flow passage 43 may include a switchback circuit that includes: a first segment that extends radially outward from a first end positioned near the inboard end of the airfoil 25 to a second end positioned near the outboard end of the airfoil 25; a second segment that extends radially inward from a first end positioned near the outboard end of the airfoil 25 to a second end positioned near the inboard end of the airfoil 25; and a crossover passage 47 that, near the outboard end of the airfoil 25, fluidly connects the second end of the first segment to the first end of the second segment. Given this configuration, the first and second segments of the pressure side near-wall flow passage 43 share a common, partitioning wall that may be configured so to maintain a fixed spaced relation between the two passages.

The suction side near-wall flow passage 44 may be similarly formed. That is, the suction side near-wall flow passage 44 may include axially-stacked and parallel first and second flow passages. As indicated, each of these flow passages 44 are defined on one side by the suction side outer wall 27 of the airfoil 25. The suction side near-wall flow passage 44 may include a switchback circuit that includes: a first segment that extends radially outward from a first end positioned near the inboard end of the airfoil 25 to a second end positioned near the outboard end of the airfoil 25; a second segment that extends radially inward from a first end positioned near the outboard end of the airfoil 25 to a second end positioned near the inboard end of the airfoil 25; and a crossover passage 47 that, near the outboard end of the airfoil 25, fluidly connects the second end of the first segment to the first end of the second segment. In this arrangement, the first and second segments of the suction side near-wall flow passage 44 share a common, partitioning wall that may be configured so to maintain a fixed spaced relation between the two passages.

It will be appreciated that the cooling passages of this type of configuration typically are formed with interconnecting rib-like structural members (hereinafter “ribs”). Such ribs 60 may be divided into two groups depending on their orientation and length. A first type, a camber line rib 62, is typically a lengthy rib that extends in parallel or approximately parallel to the camber line of the airfoil 25. (The camber line of the airfoil 25 is a reference line stretching from the leading edge 28 to the trailing edge 29 that connects the midpoints between the pressure side outer wall 26 and the suction side outer wall 27.) The second type of rib is referred to herein as a traverse rib 66. Traverse ribs 66 are the shorter ribs that are shown connecting the outer walls 26, 27 and the camber line ribs 62. The partitioning wall between the first and second segments of the pressure side near-wall flow passage 43 may be a traverse rib 66 that connects the pressure side outer wall 26 to a camber line rib 62. Similarly, the partitioning wall between the first and second segments of the suction side near-wall flow passage 44 may be a traverse rib 66 that connects the suction side outer wall 27 to a camber line rib 62. As mentioned, the central flow passage 40, the pressure side near-wall flow passage 43, the suction side near-wall flow passage 44, and the leading edge flow passage 42 have a position within the forward half of the airfoil 25. The forward half of the airfoil may be defined with reference to airfoil camber line. That is, according to embodiments of the present invention, the forward half of the airfoil may be defined as the region between the leading edge 28 and the midpoint of the airfoil camber line.

The upstream end of the pressure side near-wall flow passage 43 (i.e., the inboard end of the first segment) may be connected to a coolant supply feed 45 formed through the root 21 of the turbine blade 16. The upstream end of the suction side near-wall flow passage 44 (i.e., the inboard end of the first segment) also may be connected to the coolant supply feed 45. As stated, the first segment of both the pressure side near-wall flow passage 43 and the suction side near-wall flow passage 44 may extend from the connection made with the supply feed 45 in an outward radial direction and connect, respectively, to one of the crossover passage 47 formed near the outer tip 31 of the airfoil 25. It will be appreciated that the crossover passages 47 fluidly connect the two segments of each of the two near-wall flow passages 43, 44. Then, from the crossover passage 47, the second segment of both the pressure side near-wall flow passage 43 and the suction side near-wall flow passage 44 may extend in an inward radial direction toward the ports 46 that are positioned at the inboard end of the airfoil 25. It will be appreciated that, formed in this manner, the pressure side near-wall flow passage 43 and the suction side near-wall flow passage 44 provide an axially-stacked two-pass serpentine circuit, with the two-pass serpentine circuit of each making a 180 degree turn near the outboard tip 31 of the airfoil 25.

In an alternative embodiment, a tip flow passage (not shown) may be included near the tip 31 of the airfoil 25. The tip flow passage may connect to and be supplied by the crossover passages 47 of either of the pressure side near-wall flow passage 43 and the suction side near-wall flow passage 44. In a preferred embodiment, the tip flow passage extends parallel to the tip 31 of the airfoil 25 toward the aft portions of the blade, and may have a winding or sinusoidal configuration.

In operation, according to a preferred embodiment of the present invention, the cooling circuit described herein may operate to introduce a fresh supply of coolant via a supply feed through the root to the upstream ends of leading edge near-wall flow passages 43, 44, which may be positioned on the pressure and suction sides 26, 27 of the airfoil 25. Each of these near-wall flow passages 43, 44 may have a serpentine form that first directs the fresh coolant through an outboard segment that extends the length of the airfoil 25. The coolant may then make an approximate 180° turn and then be directed by an inboard segment that carries the coolant back to the inboard end of the airfoil 25. At this downstream end of the near-wall flow passages, ports 46 may be positioned that combine the two flows in a central flow passage 40. The central flow passage 40 acts as a plenum by which a number of radially spaced impingement connectors 48 are supplied coolant. The impingement connectors 48 may fluidly deliver coolant from the central flow passage 40 to the leading edge flow passage 42. The impingement connectors 40 are configured to deliver an impinged flow of coolant against the walls that form the leading edge of the airfoil 25. From the leading edge flow passage 42, the coolant may be expelled through surface outlets 52, which may arranged to provide film cooling to targeted surface areas of the airfoil 25.

It will be appreciated that, configured in this manner, the cooling configuration of the present invention introduces a relatively fresh supply of coolant to serpentine near-wall flow passages 43, 44 formed near the leading edge 28 of the airfoil 25. By circulating a coolant first along the outer walls 26, 27 and then combining the flows in the central flow passage 40 at a downstream location and after each has absorbed heat, the present invention reduces the temperature differential that typically occurs between the internal structure and the external walls of the airfoil 25. This will advantageously reduce the stresses that typically arise due to the unbalanced thermal expansion that high temperature differentials cause. Additionally, the present invention allows for coolant “pre-heating” such that the total cooling flow requirement is less than direct feeding the leading edge with fresh coolant. These advantages are achieved even though the cooling circuit of the present invention remains relatively simple. It will be appreciated that the simplified circuit of the present invention minimizes the pressure losses and backflow issues that come with circuits having more serpentine or torturous paths. Further, it will be appreciated that the present invention provides a configuration that is conveniently tuned. Specifically, given the separated pressure side and suction side near wall flow passages 43, 44, the configuration allows convenient adjustment of the level of coolant directed toward each side of the airfoil 25. This is advantageous given that the forward portions of an airfoil 25 have very different heat load requirements from pressure side outer wall 26 to suction side outer wall 27. As such, significant flow savings may be achieved by locally tuning convective heat transfer on each side.

As one of ordinary skill in the art will appreciate, the many varying features and configurations described above in relation to the several exemplary embodiments may be further selectively applied to form the other possible embodiments of the present invention. For the sake of brevity and taking into account the abilities of one of ordinary skill in the art, all of the possible iterations is not provided or discussed in detail, though all combinations and possible embodiments embraced by the several claims below or otherwise are intended to be part of the instant application. In addition, from the above description of several exemplary embodiments of the invention, those skilled in the art will perceive improvements, changes and modifications. Such improvements, changes and modifications within the skill of the art are also intended to be covered by the appended claims. Further, it should be apparent that the foregoing relates only to the described embodiments of the present application and that numerous changes and modifications may be made herein without departing from the spirit and scope of the application as defined by the following claims and the equivalents thereof. 

We claim:
 1. A turbine blade, comprising: an airfoil having a leading edge, a trailing edge, an outboard tip, and an inboard end where the airfoil attaches to a root configured to couple the turbine blade to a disc, wherein the airfoil further includes a cooling configuration comprising a plurality of elongated flow passages for receiving and directing a coolant along a path through the airfoil, the cooling configuration comprising: a central flow passage flanked to each side by near-wall flow passages that includes a pressure side near-wall flow passage and a suction side near-wall flow passage; a first port that fluidly connects the central flow passage to an upstream portion of the pressure side near-wall flow passage; a second port that fluidly connects the central flow passage to an upstream portion of the suction side near-wall flow passage; a leading edge flow passage; and impingement connectors that fluidly connect the central flow passage to the leading edge flow passage.
 2. The turbine blade according to claim 1, wherein the leading edge flow passage is positioned in close proximity to the leading edge of the airfoil, the leading edge flow passage extending radially outward in spaced relation to the leading edge of the airfoil from a first end positioned near the inboard end of the airfoil to a second end positioned near the outboard tip of the airfoil.
 3. The turbine blade according to claim 2, wherein the impingement connectors to allow coolant to pass from the central flow passage to the leading edge flow passage and impinge on an inner surface of the wall forming the leading edge.
 4. The turbine blade according to claim 3, wherein the leading edge flow passage comprises surface outlets through which the coolant is exhausted from the turbine blade.
 5. The turbine blade according to claim 3, wherein the plurality of impingement connectors are radially spaced between the first and second end of the leading edge flow passage.
 6. The turbine blade according to claim 2, wherein the pressure side near-wall flow passage includes axially-stacked and parallel first and second flow passages, each of which have an inner wall defined by the pressure side outer wall of the airfoil.
 7. The turbine blade according to claim 6, wherein the suction side near-wall flow passage includes axially-stacked and parallel first and second flow passages, each of which have an inner wall defined by the suction side outer wall of the airfoil.
 8. The turbine blade according to claim 1, wherein the pressure side near-wall flow passage comprises a switchback circuit that includes: a first segment that extends radially outward from a first end positioned near the inboard end of the airfoil to a second end positioned near the outboard end of the airfoil; a second segment that extends radially inward from a first end positioned near the outboard end of the airfoil to a second end positioned near the inboard end of the airfoil; and a crossover passage that, near the outboard end of the airfoil, fluidly connects the second end of the first segment to the first end of the second segment.
 9. The turbine blade according to claim 8, wherein the suction side near-wall flow passage comprises a switchback circuit that includes: a first segment that extends radially outward from a first end positioned near the inboard end of the airfoil to a second end positioned near the outboard end of the airfoil; a second segment that extends radially inward from a first end positioned near the outboard end of the airfoil to a second end positioned near the inboard end of the airfoil; and a crossover passage that, near the outboard end of the airfoil, fluidly connects the second end of the first segment to the first end of the second segment.
 10. The turbine blade according to claim 9, wherein the first and second segments of the pressure side near-wall flow passage share a common, partitioning wall that is configured so to maintain a fixed spaced relation therebetween; and wherein the first and second segments of the suction side near-wall flow passage share a common, partitioning wall that is configured so to maintain a fixed spaced relation therebetween; further comprising a sinusoidal tip flow passage that connects to at least one of the crossover passages.
 11. The turbine blade according to claim 9, wherein the first and second segments of the pressure side near-wall flow passage are partitioned by a traverse rib that connects the pressure side outer wall to a camber line rib; and wherein the first and second segments of the suction side near-wall flow passage are partitioned by a traverse rib that connects the suction side outer wall to a camber line rib.
 12. The turbine blade according to claim 9, wherein the first port is disposed near the second end of the second segment of the pressure side near-wall flow passage, and the second port is disposed near the second end of the second segment of the pressure side near-wall flow passage.
 13. The turbine blade according to claim 9, wherein the first end of the first segment of the pressure side near-wall flow passage comprises a connection to a coolant feed passage formed through the root of the turbine blade; and wherein the first end of the first segment of the suction side near-wall flow passage comprises a connection to a coolant feed passage formed through the root of the turbine blade.
 14. The turbine blade according to claim 1, wherein the turbine blade comprises a turbine rotor blade, and wherein the cooling configuration comprises a position near the leading edge of the airfoil.
 15. The turbine blade according to claim 1, wherein the central flow passage, the pressure side near-wall flow passage, the suction side near-wall flow passage, and the leading edge flow passage are disposed between the leading edge of the airfoil and a midpoint of a camber line of the airfoil.
 16. A turbine blade comprising an airfoil defined by a concave shaped pressure side outer wall and a convex shaped suction side outer wall that connect along leading and trailing edges and, therebetween, form a radially extending chamber for receiving the flow of a coolant, wherein the chamber includes a cooling configuration having: three laterally stacked flow passages positioned between the pressure side outer wall and the suction side outer wall: a pressure side near-wall flow passage disposed adjacent to the pressure side outer wall; a suction side near-wall flow passage disposed adjacent to the suction side outer wall; and a central plenum disposed between the pressure side near-wall flow passage and the suction side near-wall flow passage; and a leading edge flow passage that is positioned in close proximity and parallel to the leading edge of the airfoil; wherein ports fluidly connect the central flow passage to a downstream portion of the pressure side near-wall flow passage and a downstream portion of the suction side near-wall flow passage; and wherein impingement connectors fluidly connect the central flow passage to the leading edge flow passage.
 17. The turbine blade according to claim 16, wherein the pressure side near-wall flow passage comprises an axially-stacked two-pass serpentine circuit, wherein each pass includes an inner wall that makes contact with the pressure side outer wall; and wherein the suction side near-wall flow passage comprises an axially-stacked two-pass serpentine circuit, wherein each pass includes an inner wall that makes contact with the suction side outer wall.
 18. The turbine blade according to claim 17, wherein the two-pass serpentine circuit of the pressure side near-wall flow passage comprises a 180 degree turn near an outboard tip of the airfoil; and wherein the two-pass serpentine circuit of the suction side near-wall flow passage comprises a 180 degree turn near the outboard tip of the airfoil.
 19. The turbine blade according to claim 18, wherein each of an upstream end and the downstream portion of the two-pass serpentine circuit of the pressure side near-wall flow passage comprises a position near an inboard end of the airfoil; and wherein each of an upstream end and the downstream portion of the two-pass serpentine circuit of the suction side near-wall flow passage comprises a position near the inboard end of the airfoil.
 20. The turbine blade according to claim 16, further comprising a plurality of exhaust orifices in the leading edge flow passage for exhausting the coolant onto an exterior surface of the airfoil; wherein the impingement connectors are adapted to allow coolant to pass from the central flow passage to the leading edge flow passage and impinge on an inner surface of the wall forming the leading edge. 